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The method of characteristics has been linearized by
assuming that the flow field can be represented as a
basic flow field determined by nonlinearized methods
and a linearized superposed flow field that considers

An investigation has been conducted in the Langley
stability tunnel to determine the effects of horizontal
tails of various sizes and at various tail Lengths
(when located on the fuselage center line) and also
the effects of vertical location of the horizontal tail
relative to the wing on the low-speed static longitudi-
nal stability and on the steady-state rotary damping
in pitch for a complete-model configuration. The
wing and tail surfaces had the quarter-chard lines
swept back 45o and had aspect ratios of 4.

A summary of methods for making dynamic Lateral
stability and response calculations and for estimating
the aerodynamic stability derivatives required for
use in these calculations is presented. The pro-
ceases of performing calculations of the time his-
tories of lateral motions, of the period and damping
of these motions, and of the lateral stability bound-
aries are presented as a series of simple straight-
forward steps. Existing methods for estimating the
stability derivatives are summarized and, in some
cases, simple new empirical formulas are pre-
sented. Detailed estimation methods are presented
for low-subsonic-speed conditions but only a brief
discussion and a list of references are given for
transonic- and supersonic-speed conditions.
NACA Rept. 1102

small changes in boundary conditions. The method
has been applied to two-dimensional rotational flow,
to calculations of axially symmetric flow, to slender
bodies without symmetry, and to wing problems.

A supersonic Inlet with supersonic deceleration of
the flow entirely outside of the Inlet is considered.
A particular arrangement with fixed geometry having
a central body with a circular annular intake is
analyzed, and it is shown theoretically that this ar-
rangement gives high pressure recovery for a large
range of Mach number and mass flow and, therefore,
is practical [or use on supersoruc airplanes and mis-
siles. Experlmental results confirm gthe theoreti-
cal analysis give pressure recoverieg~p~!peT)
from 95 percent for Mach number. nt~~E~n
for Mach number 2.00. Thesq, ipsts werebgg i
m 96nally presented In a classic a cument of the NaE

Low-speed kinetic friction studies show that for
molybdenum disulfide MoS2 lubricated steel surfaces
coefficients of friction were greater at high humidity
than with dry air. MoS2 powder did not adhere to
steel surfaces at high humidity, and as a result,
metallic contact was greater and friction and wear
Increased. Increased wear with greater humidity
may have been caused by corrosion of steel speci-
mens by acids formed on contact of moisture with
MoS2. Variations in shear area resulted from
changes in humidity, slider geometry, surface finish,
and method of film application. Larger shear areas
caused greater friction when MoS2 filled the surface
interstices and was sheared.

Tests under combined axial load and Lathra 4ressure
were made of sandwich panels with simply eqipagcyked
loaded edges and free unloaded edges to determine
the strength of panels of various thicknesses and to
compare the results with computed values. The
theory, derived in this paper, is based on the theory
for buckling of simply supported sandwich column
and was conservative in predicting larger strains or
deflections than those measured. Agreement
between computed maxmum loads and exprimental
failing loads was within 9 percent.

NACA TN 3093

EFFECT OF TYPE OF POROUS SURFACE AND
SUCTION VELOCITY DISTRIBUTION ON THEE
CHARACTERISTICS OF A 10. 5-PERCENT-THICKE
AIRFOIL WITH AREA SUCTION. Robert E.
Dannenberg and James A. Weiberg. December
1953. 59p. diagrs., photos., 5 tabs. (NACA
TN 3093)

Results are presented of an investigation of a two-
dimensional, 10. 51-percent-thick symmetrical air-
foll with area suction near the leading edge. Lift
and suction-flow characteristics were determined
wsth different porous surfaces (perforated plates
and sintered steel) for various suction velocity dis-
Iributions. The flow requirements were ascertained
over a range of free-stream velocities. The maxi-
mum lift was independent of the surface of the
materials tested.

The effects of diffusion processes on the smoking
tendencies of eight laminar diffnaton flames were in-
vestigated by varying the rate and concentration of
air and oxygen supplied to the flame. Increasing the
rate at which air was supplied permitted rather Lim-
ited increases in the smoke-free burning rate. In-
creasing the concentration of oxygen from 21 to 45
percent increased the smoke-free fuel flow for
pentene-1, neopentane, isobutane, ethylene, and
n-butane, but decreased the smoke-free burning rate
for butene-1, cyclopropane, and propene in the range
of 25 to 40 percent oxygen. Increasing the flame
temperature caused the smoke-free fuel flow to in-
crease except for butene-1, cyclopropane, and pro-
pene, where initial rates were lower. The varia-
tions in smoke formation with oxygen enrichment
were tentatively explained by considering the effect
of temperature on the initial decomposition reactions
of the fuel molecule.

An analysis is made of the transient heat-conduction
effects an three simple semi-infinite bodies: the in-
sulated flat plate, the conical shell, and the slender
solid cone. The bodies are assumed to have constant
initial temperatures and at zero time to begin to
move at a constant speed and zero angle of attack
through a homogeneous atmosphere. The heat input
is taken as that through a laminar boundary layer.
Radiation heat transfers and transverse temperature
gradients are assumed to be zero. The appropriate
heat-conduction equations are solved by an iteration
method, the zeroeth order terms describing the situa-
tion in the limit of small time. The method is pre-
sented and the solutions are calculated to three
orders which are sufficient to give reasonably
accurate results when the forward edge has attanlned
one-half the total temperature rise (nose half-rise
time). Flight Mach number and air properties occur
as parameters an the result. Approximate extpree-
sions for the extent of the conduction region and nose
half-rise times as functions of the parameters of the
problem are presented.

Results are shown of an investigation of tunnel-wall
Interference in a twro-dimensional-flow, rectangular,
closed-throat wmd tunnel through a Mach number
range from 0.3 to 0.9 and a corresponding Reynolds
number range from 0.9 x 106 to 1.8 x 106s. For
ratios of airfoil chord to tunnel height below approi-
mately 0.15, correction of aerodynamic data for wall
interference by the small-perturbation theory of
NACA Rep. 782 was found to yield results in satis-
factory agreement with essentially interference-free
data.

One-dimensional, steady-state, compressible, vis-
cous flow relations are presented which permit the
determination of flow conditions at any radial posi-
tion in a ducted helicopter blade. The relations are
required for estimating the performance of proposed
helicopter jet-propulsion systems which Involve duct-
ing air or gases through the blade from root to tip.
A limited number of calculations over a wide range
of helicopter operating conditions and relative duct
sizes are also presented. The choking" problem in
the straight duct is discussed,

Flight measurements hav~e been made of the phugoid
motion of the Hoverily Mk. i nellcopter, following an
arbitrary longitudinal displacement of the control,
the latter being returned to its Initial position and
held fixed. The tests were done throughout the
speed range for power-on conditions and in autorota-
tion for various center-of -gravary positions and for
forward and backward Initial displacement of the
stick.
N-27231'

An investigation is made into the characteristics of
a freely suspended flexible sheet as a shock absorber
replacing the conventional undercarriage, particular
attention being given to the inertia of the sheet. It
is found that when an aircraft as dropped vertically
an to the sheet the retarding force is first produced
by the inertia of the sheet Itself, and not until later
in the descent by the reactions from the side sup-
ports of the sheet. By careful adjustments of the
mass and tension of the sheet "retardation effi-
ciencies" exceedi~ng 80 percent can be achieved.
The effect of the aircraft having a forward component
of velocity Increases the contribution of sheet mo-
mentum. For reasonably practical landing speeds
and sheet dimensions, virtually the whole of the mlo-
mentum of descent is absorbed by sheet inertia.
Under such candations still higher retardation efft-

clencies are obtainable and, with a suitable design of
aircraft keel, rebound may be entirely eliminated.

This report describes tests carried out on the 30
percent Griffith symmetrical airfoil with continuous
suction applied through a porous capping fitted over
the front 15 percent of the upper surface. Through-
out the range of incidence covered an the exp~eri-
ments, distributed suction was found to decrease the
slot suction necessary to prevent separation, espe-
cially when the distributed suction caused rearward
movement of the transition position. The profile
drag of the airfoil was measured, and estimates
were made of the equivalent drag coefficients for the
work done by the suction pumps. Assuming no
Losses additional to those in the boundary layer, it
was found that the effect of distributed suction was to
reduce slightly the overall drag of the alrfoll.
Measurements of the velocaly within the boundary
layer were made at various chordwise positions on
the porous surface; the profiles recorded were very

characteristic spread of turbulent flow in the wake
ofa sanlu panrtce on the surf a e waentuch r du ed

The problem of the estimalaon of the aerodynamic
forces acting on two-dimensional atriodls ascallating
at mean Incidences below the stall Is considered A
method of calculation is suggested which makes use
of the steady motion characteristics of the airfoil.
At low frequencies, good agreements withl the meas-
ured aerodynamic derivatives should be obtained as
the method is such that it giv~es the correct values at
zero frequency. A comparison between the esti-
mated and measured values of the pllchlng-moment
derivatives for a particular airfoil Is made, and this
shows that the method suggested gives better agree-
ment with experiment than the usual vortex-sheet
theory.

The validity and accuracy of methods of determining
correasions to the measured velocity in a wind tunnel
Lo compensate for the constraining effect of the walls
are reviewed following recent experimental evidence
from the R. A. E. 10- by 7-foot subsonic wind tunnel.
It is concluded that such corrections, commonly
known as "blockage" corrections, can be success-
fully applied at Mach numbers up to 0.96 but some
modifications are necessary to the formulas at
present in use. Formulas for the calculation of he
longitudanal distribution of blockage increment due
to any model, necessary to check the validity of the
method in particular cases, are presented in a form
which, it is hoped, will facilitate their use an any 10
by 7 wind tunnel. Formulas for the corresponding
wall velocity increments, used to check the accu-
racy of the method by comparison with measured
wall pressures, are also given.

t~ur ane adds nr de rti othe ba es hav cn-
ventional sections including both reaction and Im-
pulse designs. The two-dimenslonal performance
over wide ranges of incidence at Mach numbers up to
1.0 is discussed, special importance being attached
to the effects of compressibiliy. It is shown that
the effec teom ceasin othe dgreenofr at n is o

:'ale ":1dec erane calhi Mach numbers a on
loss watin these particular blade designs, while the
cos-1 throat pitch rule is found to be approximately
true only at hagh Mach numbers.

The inefficient pressure recovery of present-day
supersomec wind tunnels, which leads to nigh costs
of plant installation and operation, is discussed and
methods of improvement suggested. In particular.
the diffuser system, where most of the losses occur,
is studied in detail; the improvement to be expected
in the pressure recovery by the use of convergent-
divergent types ia explained and methods of over-
coming the necessity for high starting powers with
this arrangement are presented. Diffuser experi-
ments based on recent investigations info breakaway
phenomena in supersonic flow are described which
result in a considerable improvement of pressure
recovery. A deceleration from M =2.48 at the
working section to M =1.42 at the dlfuser throat
was obtained using a variable diffuser throat.

The theory of the flow through a throat near sonic
velocity is developed, and la followed by a discus-
sion of the conventional method of designing super-
sonic nozzles using the method of characteristics.
A method of improving the Mach number distribution
of the nozzle using the experimental results is devel-
oped. The nozzles designed were tested in a 3-inch
square wind tunnel in which the Mach number distri-
bution was obtained by shaping the top wall of the
working section. The Mach number distribution
along the bottom wall was determined from the pree-
sures measured by a series of static-pressure holes
along the wall. Considerable difficulty was found in
improving the distribution; this was considered to be
due to the discontinuity in curvature at the point of
inflexion and the influence on the boundary layer of
the sudden relaxation of the pressure gradient along
the wall. An alternative method of design was de-
veloped which avoided this discontinuity in curva-
ture, and considerably better results were obtained
when attempts were made to improve the experi-
mental Mach number distribution.

A rigorous theory has been developed for deter-
mining the stresses and displacements in a sheet
reinforced by strangers and ribs which are not at
right angles to the stringers. The solution of many
problems of practical importance has been facilitated
by the introduction of a stress function. The theory
has been applied to a cylinder of rectangular section
stiffened with such skew ribs (a simplified represen-
ration of a swept wing). It is shown that there are
axes about which applied moments produce pure
twist or pure curvature of the cylinder. There are
simple formulae for determining these axes and the
relationships between twist and curvature and the
applied moments.

The report summarizes the more practical aspects
of the results of a long-term investigation of the
basic physical and chemical properties of polymethyl
methacrylate ("Perspex" type) plastic. Thermal,
elastic, craving, solvent absorption, and mecharucal
properties are included and the effect of these on the
service efficiency of a plastic structure is described.
Experimental evidence is given concerning the essen-
tial role of tensile arress and absorbed solvent in
causing crazing and recommendations concerning
means to reduce or avoid the incidence of crazing
are included. The basic thermal properties are
compared with those of metals and the dangers of
differential expansion in combined metal-plastic
structures are noted, together with the serious ef -
fects of chilling of plastic structures during the
"hot-forming" operation. Details are gaven of ap-
propriate heat treatments designed to remove castinB
and workshop strains without causing distortion.

Experimental work is now being done to establish a
basis for the solution of whirling problems on tur-
bine and contrarotating shaft systems in the design
stage. This report is concerned primarily with the
degree of accuracy to be expected from experiments
on models. In experiments here described results
are obtained for a simple cantilever system which
are in close agreement with theory. With more
complicated systems the error is somewhat greater
owing to practical effects not covered by theory,
though still acceptable for most design purposes.

Part I of the report gives an account of experiments
made with the National Phyascal Laboratory smoke
ge"rto ".i obaingnasmo~ke pattrs ann ath ake of an

form motion relative to (a) an infinite row of equally
spaced two-dimensional discrete vortices of alter-
nate sign and (b) an infinite twoe-dimensional vortex
sheet with sinusoidal distribution of strength. Com-
n risn wtel certain of the smoke patterns discussed
treated very closely as a system of discrete vortices,
and this is supported by the consideration that the
elements of a continuous vortex sheet would in gen-
eral be subject to normal induced velocities, which
would tend to break the sheet up. The tendency for
this to occur would be greater for the higher values
of ra, wince the variation of vorticity with distance
along the wake would be greater. The induced ve-
locities for a uniform infinite vortex sheet would be
zero.

The investigations described in this report were
undertaken for the Aeronautical Research Committee
an order to study the effect of porosity on the sta-
bility of a parachute. Experiments were conducted
to obtain data which could be applied in the usual
way to the standard equations of motion.

A simplify sed panel model is described, together
with a number of assumplions about the mode of its
buckling. The approach to the calculation of the
buckling stress as by splitting the panel into a num-
ber of flat plates and Ireating these by the ordinary
plate theory. Use of the boundary conditions be-
tween these plates leads to a reltaion between the
buckling stress and the variables of the panel geom-
etry. The results thus obtained are compared with
two sets of recent experlmental work, and an appen-
dix Is included to show the effect of initial panel ir-
regularities on the experimental determination of
buckling stresses.

Al Ithe plates to tedkwereh3e5inche nongb nd nom -
supports was varied between 35 and 120 times its
thickness. Both clad (D.T.D. 546) andunclad
(D.T.D. 646) maternal were tested. Three types of
edge support were used: rowrs of steel balls an vee-
goove blocks, i tendedt Irsmianm 0snseded con

Intended to smatate clamp-edged conditions; and a
single type of stringer used in previous panel tests.
Measurements were made of the plate load and mean
strain, and of the shape of the skin buckles. The
test techmque is discussed and the experimental re-
suits compared with theory.

straight winge being dealt with as special ssee
Complete detailed solutions are given for wings Ath~
biconvex parabolic profile, and the problems in-
volving arbitrary profiles are investigated in a more
general way. The velocity distribution in the
regions of kinks and tips differing considerably
from that on an infinite sheared wing, some local
positive or negative drag arises in these regions.
The drag drstrlbution Is calculated for the case of a
biconvex parabolic profile, and it is found that,
while the total resultant drag must obviously be nil
for a fimite wing, at may differ from zero for infinite
or semi-infinite wings. This ''potential drag is
generally small.

Schlleren tests on a series of conventional turbine
cascades have shown that the variations in perform-
ance at high speed can be accounted for by shock-
wave and boundary-layer interaction. The rise in
Ioss coefficient sometimes encountered at outlet
Mach numbers of 0.6 to 0.8 is shown to be due to the
formation of a A shock series on the upper surface
of the blade, the subsequent fall in loss coefficient
and increase in deflection as the outlet Mach number
rises to unity being caused by the formation of a
shock system at outlet which forces the separated
part of the boundary layer back on to the blade sur-
face. It isshown that a shock series may form
on a boundary layer which is apparently turbulent,
This has not been observed before.

Owang to the abrupt change in shear stress at loading
sections of beams there is a concentration of direct
stress in the outer fibers of the beam near the load-
Ing section. A method of calculating this concentra-
tion is described. The highest stress concentrations
occur in short deep beams and are greater for wood-
en than metal beams. The method is applied to the
spars of two wooden aircraft and stress concentra-
tions 1.06 and I.4 are found at the fuselage attach-
ments. Strain measurements were made at posi-
thons on a wooden beam under load and the theorett-
cal predictions verified.
N-27259*

This report contains information of all the revelant
tables known to the author, which can be obtamned by
workers outside the establishment of origin. The
purpose of the index is to make available to workers
in the field of compressible flow a reference from
which they may trace a tabulation of any function
they require, if it exists.
N-27254,

Measurements of the thrust, torque, and flapping
angle for a 12-foot diameter rotor over a range of
blade angle, shaft inclination, and tip-speed ratio
have been made to give information on the validity of
the standard rotor theory and of the effect of stalling
on the retreating blade. Good agreement with the
theory was obtained over the normal operating range'
using airfoil characteristics determined from the
measurements an the static thrust condition. Stalling
was found to be progressive in character showing
first by an increase in torque and flapping angle and
lae ya fall in thrust, as compared with the calcu-

Two simple means for establishing a relation be-
tween a pair of oscillation problems are briefly dis-
cussed. In the first, the displacements are con-
nected by use of a differential operator. The set of
natural frequencies rs identical for the two problems
and results of interest are obtained when the trans-
formed boundary conditions can be physically inter.
preted. In this manner, it is shown, for example,
that a flywheel on a uniform shaft can be transformed
into a flexible coupling and a mass carried on a uni-
form beam into a flexible hinge. In the second, the
connection is established by use of the concept of
mechanical admittance. Here the frequency equa-
tions are simply related but the frequencies are not.
N-27256

This report is a continuation of an earlier one and
puts forward several new solutions of the problems
of velocity distribution on finite or semi-infinite un-
tapered wings, at zero incidence. The solutions
are based on the first order method of sources and
sinks, which is shown to be sufficiently accurate to
deal with problems involving tips or kinks. The
fundamental case considered le that of a semi-
infinite sheared wing, and the theory is built up to
embrace finite sheared and sweptback wings,

NACA
RESEARCH ABSTRACTS NO.S55

This note explains an improved numerical method of
evaluating the contributions to the downwash at
moderate or large spanwise distances from a vortex
lattice. By allowing freedom of choice of the chord-
wise positions of the discrete vortices of the lattice,
it is possiljle to select three definite chordwise posi-
tions and strengths of vonrtles at each of these posi-
tions dependent on the chordwise pressure distribu-
tion, so as to determine the downwash with good
accuracy for three particular pressure distributions

proportional to cot Z 9, sun 9, and sin 20. The
corresponding chordwise loading factors have also
been evaluated for deflected flaps.

The paper describes and applies exact methods of
calculating the Incompressible flow about thick alr-
foils of general shape in a free stream, and about
symmetrical airfoils between channel walls. One of
these methods is extended to an approximate treal-
ment of subsonic compressible flow by making use
of von Ka'rmin's transformation.

Knowledge of the variation close behind a traiLing
edge of the wake displacement thickness

dfj = 1- y is necessary In calculations of
the circulation round an airfoil. Examination of
data now available reveals that in the wake: (1)
velocity profiles on either side of the hine of mini-
mum velocity may be derived from the corresponding
trailing-edge boundary-layer profiles by change of
scale of each coordinate; (2) the velocity defect at
corresponding points follows a universal recovery
law ~ ~ ~ o h o -
lawof he ormK(x zol even close to the
trailing edge. An immediate consequence of these
two empirical properties is a simple relation for the
form parameter H = 5*, 9 at points in the wake In
terms of tralling-edge values. In conjunction w~ith
the momentum equation, this makes 6* determi-
nate. Agreement with experiment is very satls-
factory.
N-27264*

Aeronautical Research Council (Gt. Brit.)
NOTE ON THE LIFT SLOPE, AND SOME OTHER
PROPERTIES, OF DELTA AND SWEPT-BACK
WINGS. E. F. Relf. 1953. 9p. diagrs., tab.
(ARC CP 127)

In studying and comparing various theories for the
determination of the distribution of loading on wings,
Garner has given values for the hit slope of several
families of sweptback and delta wings deduced from
several different hiting-surface theories. In fIg-
ure 8 of reference 1, Garner has plotted these lift
slopes as functions of the aspect ratio A, for dif-
ferent values of the angle of sweep. It occurred to
the writer to try plotting the ratio of the l1If slope to
that for elliptic loading instead of the laft slope utself,
and when this was done it was noticed that the above
ratio was very nearly independent of aspect ratio A,
and gave a unique curve for all the available results
when plotted against sweepback angle A. It has
been possible to make some comparisons of theoret-
ical deductions drawn from the above with measure-
ments made in the C. A. T. on four delta wings, one
swept wing, a tapered swept wing on a body, and two
untapered swept wings on a body at high Reynolds
numbers, where one could expect a close approxi-
mation to potential theory. While analyzing the
C.A. T. results, values of the quantity K in the

formula CD =CDo Aa CL2 were collected and
studied. Lastly, a study was made of scale effects
In the C. A. T. tests.

N-27265'

Aeronautical Research Council (Gt. Brit.)i
METHOD FOR THE DETERMINATION OF THE
PRESSURE DISTRIBUTION OVER A FINITE THIN
WING AT STEADY LOW SPEED. G. J. Hancock.
1953. I Ip. diagrs. (ARC CP 128)

For any given pressure dlstrlbutlon across a finite
thin wing at lowr speed, the wing surface can be ob-
tained by direct double integration. Therefore, the
pressure distribution across a given wing surface
may be obtained by the superposition of a number of
solutions rn which the wing surface is known for a
prescribed pressure distribution. The method has
been applied for the determination of the pressure
distribution across a thin uncambered delta wing.

N-27266~'

National Gas Ttrrbine Establishment (Gt. Brt. )
AN ANALYSIS OF THE AIR FLOW THROUGH THE
NOZZLE BLADES OF A SINGLE STAGE TURBINE.
I. H.Johnston. 1953. 18p, diagrs. IARC CP 131)

This memorandum presents the results of detailed
traverses made on three of the nozzle assemblies
designed for a single stage experizriental turbine.
The effects of pitch,'chord ratio on gas outlet angle
and total head loss are recorded and discussed in
the light of corresponding work published elsewhere.
The value of pitch, chord ratro giving minimum total
head loss is found to compare well with the optimum
pitching given by two-dimensional results obtained
from cascade tests on blades of a similar nature.
N-27267*

NACA
RESEARCH ABSTRACTS NO. 55
Four species of wood were evaluated as tatheir use-
fulness in the manufacture of plywood. Da~a ag
their processing, drying, gluing qualittee, i a
pressure compression properties is given.

This memorandum presents the results of a series
of low speed tests on six stages of a medium stagger
free vortex design of axial compressor blading, in
which the stagger of the stator blades was varied
over a wide range while the rotor blade stagger re-
mained at its design figure. It is shown that effi-
clencies an excess of 85 percent were achieved over
a range of stator blade stagger from -500 to *100,
compared with the design figure of -25. 4o. This
performance augurs well for the Improvement of the
performance of axial flow compressors away from
their design point by the method of altering the stator
blade stagger in some of the stages.
N-27268'

Emmons has given a relaxation method of dealing
with shock waves when the compressible stream
function is the dependent variable. This paper
briefly outlines a procedure to adopt when log
(q = velocity magnitude) le taken as the dependent
variable. A method of allowing for the presence of
vorticity behind the shock wave is also given.
N-27287'

An investigation on the effects of short periodl heat-
ing of aluminum alloys in the range 1200 C -2000 C
was made. Curves are presented showing the of -
fects of heating on the 0. I percent proof stress and
ultimate stress of the material. A further investi-
gation was made on the possibility of recovery of
properties taking place either by prolonging the low-
temperature storing time or by leaving the material
for a long time at room temperature after shoving.

Further measurements were made in the working
section of this 4-foot by 3-foot wind tunnel. The
influence of the number of screens in the bulge on
the turbulence in the working section was studied,
and the intensity and scale of turbulence was
measured at the end of the second diffuser and at
various places downstream and rn the bulge.

Aor end the es Mnse Lc on of Ih fis c r ece
of maximum section lift on sweptback wings is pre-
sented. The method utilizes simplified lifting-
surface theory, two-dimensional data, and simple
sweep theory. The procedure is applied to a swept-
wing model with and without trailing-edge split flaps
and with several wing modifications. The predicted
wing lift coefficients for the onset of stall are com-
pared with measured values at which marked changes
in force and moment characteristics occurred.

The effect of interference on the wave drag of a
combination of bodies of revolution at zero angle of
attack at supersonic speeds is investigated. Numeri-
cal calculations of the drag change obtainable from
interference are carried out and curves are drawn
for cases of two bodies of identical fineness ratto but
with lengths inl the ratio 2 to 1. Also considered is
the special case of a three body combination with
bilateral symmetry, for which it is found that the
total wave drag can be 35 percent less than the total
wave drag of the same bodies without mteractlon.

Free-flight drag measurements of 300, 600, and 90o
cone cylinders, with cylinder fineness ration of 1.2,
at Mach numbers between 1.5 and 8. 2, with Reynolds
numbers in the order of 1 million, are presented. It
is concluded that the Taylor and Maccoll theory for
wave drag of cones is accurate for 600 cones at
Mach numbers from 2 to 8 in the absence of gaseous
imperfections. The base drag of the 600 cone cylin-
der is calculated from the experimental results at
Mach numbers 2.0 to 4.5. Discontinuities along a
streamline in the flow about and behind the models
are observed and explained.

A description of the instruments used in the Ames
supersonic free-flight wind tunnel to obtain a time-
distance record of free-flying models over a 15-foot
flight path is presented. The metruments include a
chronograph and shadowgraph stations. The meas-
uremeintcshareC ccuratet within 0.1 mi rosecond and

Research models are fired from a gun through the
wind-tunnel test section in a direction opposite to the
air stream which has a moderate supersonic Mach
number. A wide range of Mach numbers, from low
supersonic speeds up to M = 10, can be obtained.
The equipment and test techniques are described.
The methods used to measure drag, Llft-curve slope,
center of pressure, and damping in roll are given.
The imperfections in the air stream and their effect
on model tests are discussed.

NACA RM A52D22

THE FORCES AND PRESSURE DISTRIBUTION AT
SUBSONIC SPEEDS ON A CAMBERED AND TWISTED
WING HAVING 450 OF SWEEPBACK, Ab] ASPECT
RATIO OF 3, AND A TAPER RATIO OF 0.5.
Frederlck W. Boltz and Carl D. Kolbe. July 1952.
166p. dlagrs., 22 tans. (NACA RM A52D22)

Lift, drag, pitching-moment, and pressure data for a
model of a 450 swleptback cambered and twisted wing
having an aspect ratio of 3 and a taper ratio of 0.5
are reported. The alrfoll sectsons were the NACA
64A410 In planes inclined 450] to the plane of symme-
try. Data are presented for Reynolds numbers from
4,000,000 to 18,000,000 at a Mach number of 0.25, for
Reynolds numbers from 41,000,000 to 8,000,000 at a
Mach number of 0.60, and for Mach numbers from
0.08 to 0.96 at a Reynolds number of 4,000,000. A
comparison of the data with those for a plane wing of
identical plan form is made. Pressure data at seven
spanwise stations on the model are presented an tabu-
lar form.

NACA
RESEARCH ABSTRACTS NO.SB

a blade rowr. Quantitative turning a corrections
due to effects of secondary flows in -rifleet coz'q-
pressor inlet guide vanes were obtained from indluced
deflections of a superimposed vortex system in con-
junction with an empirically determined correlation
factor.

The experimental performance of turbine nozzle
blades designed for a constant discharge angle was
investigated at discharge hub Mach numbers of 1.18,
1.31, and 1.41l. Flow characteristics are presented
in terms of energy losses, angle gradients, and
secondary flow effects. Blade efficiency decreased
from 0.983 to 0.978 with increasing Mach number in
the range investigated while angle variations in the
loss regions became very large, indicating poorer
blade performance than efficiency implies.
NACA RM E52D07

An analysis is presented to give a qualitative picture
of the operation of each stage in a hlgh-pressure-
ratio multistage compressor over a full range of
operating flows and speeds and to point out methods
of improving off-design performance. Single-stage
performance results have been "stacked" to form a
multistage compressor in which the design or match
point of each stage has been arbitrarily selected.
The effects of single-stage performance, stage
match pomnt, and stator-blade-angle adjustment are
considered

ANALYTICAL INVESTIGATION OF TURBINES
WITH ADJUSTABLE STATOR BLADES AND EF-
FECT OF THESE TURBINES ON JET-ENGINE
PERFORMANCE. David H. Silvern and William R.
Slivka. July 17, 1950. 51p. diagrs. (NACA
RM E50E05)

Adjustable-stator turbines are applied to turbojet
engines and probable performance is compared with
conventional engines with and without variable-area
exhaust nozzles. Variation In stator-exit angle and
e hut area wasa net exssv fo e cags of

contemporary turbanesmequipped withra 917 bl

percent in overall engine specific fuel consumption
over conventional engines and from 2 to 8.5 percent
over engines equipped with only adjustable-area
exhaust nozzles were obtained at 60-percent rated
power with adjustable-stator turbines and variable-
area exhaust nozzles. The Improvements depend
on design parameters.
NACA RM E51F25

An investigation was conducted in a 3.4- by 3.4-inch
duct to determine the characteristics of the super-
sonic flow downstream of four wire-mesh screen
nozzles with nominal design Mach numbers in the
range between I.97 and 2.58. Two types of dlsturb-
ance were observed In the flow field: a fine network
of interacting expansion and compression waves
which formed immediately downstream of the
screens and appeared to dissipate within 25 to 40
wave Intersections, and relatively strong oblique
shock waves that originated at the functions of the
screens and the walls and were reflected throughout
the length of the duct. Regions of fairly
uniform flow were found to exist. The total-
pressure loss across the screens varied from 22
percent at Mach number 1.58 to 43 percent at Mach
number 2.06.
NACA RM E51G27

Qualitative discussion is presented of general nature
of secondary flows in stationary annular cascades
with thin wall boundary layers and radial design var-
lation of circulation. Deviations from Ideal mean
outlet flows (based on blade-element performance)
exist in potential-flow region of vanes because of
comins ipsdab en-al bonas,sclis ace-
surfaces, and Irrotationallty requirement. As a con-
sequence of existence of nonuniform radial flow
across blade spacing, it may not generally be possi-
ble to obtain any arbitrarily specified design varia-
tion of turning angle along the radial height of

NACA
RESEARCH ABSTRACTS NO.55

An analysis is made of the factors affecting the
weight-flow rate per unit frontal area of centrifugal
compressors with axial-flow vaned diffusers pre-
ceded by mixed-flow vaneless sections. It is shown
that, for specified inlet conditions to the impeller
and vaned diffuser, the weight-flow rate is increased
at the expense of pressure ratio and vice versa.
Charts are presented to help the designer make a
satisfactory compromise between weight-flow rate
and pressure ratio. Some conclusions of the inves-
tigation are: (1) Prewhirl is of negligible value In
centrifugal compressors designed for high weight-
flow rates. (2) Transonic Inlet flow conditions are
desirable for high values of compressor weight-flow
rate and pressure ratio. (3) Atthe inducer tlpa
value of 60o for the inlet relative flow angle, meas-
ured from the axial direction, results in approxl-
mately maximum values of compressor weight-flow
rate per unit frontal area.
NACA RM L8HO4

AERODYNAMIC CHARACTERISTICS OF TWO ALL-
MOVABLE WINGS TESTED IN THE PRESENCE OF
A FUSELAGE AT A MACH NUMBER OF 1. 0.
D. William Conner. October 28, 1948. 20p.
diagrs., photos., tab. (NACA RM L8HO4)

Half-span models of two wings of different plan form
were tested both as all-movable surfaces and as
fixed surfaces in the presence of a half fuselage in
the Langley 9- by 12-inch supersonic blowdown
tunnel at a Mach number of 1.9. One wing had a
half-delta plan form with 600 leading-edge swee@
and was tested at a Reynolds number of 1.9 x 10 .
The other wang had a rectangular plan form modified
by an Ackeret type tap and was tested at a Reynolds
number of 1.4 x 106' .

NACA RM L8H I2

YAW CHARACTERISTICS OF A 520 SWEPTBACK
WING OF NACA 641-112 SECTION WITH A FUSE-
LAGE AND WITH LEADING-EDGE AND SPLIT
FLAPS AT REYNOLDS NUMBERS FROM 1.93 x 106
3306. gr 6.phReino .Salml November 8, 1948.

Contains results of low-speed, high Reynolds num-
her, wind-tunnel tests of a 520 sweplback wing in
yaw. The effects of Reynolds number, flap deflec _
tion, and fuselage position on the static-lateral-
stability parameters are given. Includes data at
high yaw angles and effects of fuselage position on
the sidewash characteristics in the region of a verts-
cal tail*

An investigation has been made in the Langley free-
fIught tunnel to determine the lateral stability of a
flying model equipped with a gyro stabilizing unit
which applied control in response to bank and yaw.
The results are presented in the form of time his-
tories of motions of the flying model with flicker
and hunting type of control. A systematic calibra-
Lion was made, and formulas were developed to de-
termine the response of the gyro unit to angles of
bank and yaw for various angles of cant and tilt.
NACA RM L9Bl8

A wind-tunnel investigation of a variable sweep,
complete model was conducted at angles of sweep-
back of 450, 300, 150, and Co. The investigation
included the effect of various wing cutouts, a sharp-
leading-edge wing section, a wing vane, flap deflec-
lion, and vertical location of the horizontal tail.
The data permal an estimation of the amount of
longaludinal wing translation required to compensate
for the stability changes accompanying the change in
sweep angle.
NACA RM L9EO2

THE E EFFECT OF SPAN AND DE FLECTION OF
SPLIT FLAPS AND LEADING-EDGE ROUGHNESS
ON THE LONGITUDINAL STABILITY AND GLIDING
CHARACTERISTICS OF A 42o SWEPTBACK WING
EQUIPPED WITH LEADING-EDGE FLAPS. George
L. Pratt and Thomas V. Bollech. June 21, 1949.
26p. diagrs., photo. (NACA RM L9EO2)

The effect of half-span and full-span split flaps
through a deflection range of Oo to 600 on the low-
speed longitudinal and power-off gliding character-
Ishecs of a wing with 42o sweepback at the leading
adg eu ppedewith leading-edgheef aps and tie effect
ability of the wing equipped with leading-edge flaps is
presented. The split-f lap tests we~re made at a
Reynolds number of 6.8 x 106, and the effect of
leading-edge roughness on the wing equipped with
leading-edge flaps was determined at Reynolds num-
bers of 3.0 x 106 and 4.7 x 10 .
NACA RM L9E24

INVESTIGATION OF LOW-SPEED AILERON CON-
TROL CHARACTERISTICS AT A REYNOLDS NUM-
BER OF 6,800,000 OF A WING WITH LEADING
EDGE SWEPT BACK 420 WITH AND WITHOUT
HIGH-LIFT DEVICES. Thomas V. Bollech and
George L. Pratt. July 19, 1949. 31p. diagrs.,
photo. (NACA RM L9E24)

Presents the lateral control, hinge-moment, aileron
load, and balance-chamber-pressure characteristics
of an aileron on a wing with leading edge swept back
420 with and without high-lift and stall-control de-
vices. The Investigation was carried out at a
Reynolds number of 6,800,000 for an angle-of-attack
range from -4o through the stall.

Contains results of wind-tunnel tests at a Reynolds
number of 6.0 x 106 (Mach number of 0.14) on the
effects of leading-edge and trailing-edge flaps on the
longitudinal stability characteristics of two 471.70
sweptback wings of aspect ratios 5.1 and 6.0. The
effects of roughness, fuselage interference, and
wing fences are also presented.

A survey of various types of entapults, which has
been made mn connection with the problem of accel-
erating a large (100,000 lb) car along a track to a
speed of 150 miles per hour, is given. A hydraulic
jet catapult is indicated as the best-suited among
these catapult types for the purpose intended, and
various design problems of this type are treated.
Equations are gzven for calculating the performance
of the jet and of the test ear, and consideration is
given to the physical conditions affecting the jet flow.
Design procedures are presented for the jet nozzle
and for the bucket on the car which receives the jet
and imparts thrust to the car. The expected pro-
pulsive efficiency of the jet catapult is given and the
effect of a side wind on the let trajectory is calcu-
lated.

A delta wing and a tapered sweptback wing of aspect
ratios 1. 56 and 2. 00, respectively, both of which
were cambered and twisted so as to provide a uni-
form load distribution for a supersonic flight con-
dition, were tested in combmation with a fuselage
at Reynolds numbers between 384, 000 and 1,550, 000
an order to determine the low-speed lift-drag and
static stability characteristics of such wings.

An investigation was made in the Langley 19-foot
pressure tunnel of a wing incorporating NACA
65-210 airfoil sections and having an aspect ratio
of 5.8. High-Llt and stall-control devices included
split, single, and double slotted trailing-edge flaps;
slats, extensrble-, and drooped-nose flaps. Static
longitudinal stability characteristics with midwing
fuselage and several vertical locations of a horizon-
tal tail were determined with flaps neutral and de-
flected. Most of the data are presented for a
Reynolds number of 6.5 x 106 corresponding to a
Mach number of 0.2. Lift, drag, and pitching-
moment data are presented for all configurations.
Some stall diagrams are presented.

Presents the effects of spoiler geometry and location
on the low-speed lateral-control characteristics of a
wing swept back 420 at the leading edge with and
without high-lift and stall-control devices. The
investigation was carried out at a Reynolds number
of 6.8 x 106 and through an angle-of-attack range
from -40 through the stall.

Contains results of low-speed wind-tunnel tests on
two wings of 47.70 sweepback and aspect ratios of
5.1 and 6.0. Aerodynamic data are presented for a
range of Reynolds numbers from about 1.1 x 106 to
10.0 x 106. Results show that abrupt unstable
pitching moments occur at moderate lift coefircients.
The maximum lift coefficients are about 1.20 and are
only slightly affected by Reynolds numbers, but a
considerable scale effect occurred for the lift coeffi-
cient at which the pitching moment broke. A maxi-
mum lift-drag ratio of about 27.8 was obtained with
the aspect ratio 6.0 wing.

NACA
RESEARCH ABSTRACTS NO.55

NsACA RM L51E24

LOW-SPEED INVESTIGATION OF THE EFFECTS OF
SINGLE SLOTTED AND DOUBLE SLOTTED FLAPS
ON A 47.70 SWEPTBACK-WING FUSELAGE
COMBINATION AT A REYNOLDS NUMBER OF
6.0 x 106. Ernst F. Mlollenberg and Stanley H.
Spooner. September 1951. 23p.diagrs., photos.,
3 tabs. (NACA RM L51E24l)

The effects of deflection of partial-span single and
double slotted flaps in combination with leading-edge
flaps on the longitudinal aerognamic characteristics
al a 47.70 sweptback-wfug fuselage combination are
abown. The wing had an aspect ratio of 5.1, a taper
ratiO of 0.383, and NACA 64-210 airfoil sections.
The tests were conducted in the Langley 19-foot
pressure tunnel at a Reynolds number of 6.0 x 106
and a Mach number of 0.14.

Results of pressure-distribution tests at Reynolds
numbers of 1,500,000 and 4,000,000 and of force tests
through a Reynolds number range from 1,500,000 to
4,800,000 are presented. Pressure data for
4,000,000 Reynolds number are tabulated and data for
both Reynolds numbers are presented as wing and
section force and moment characteristics. Effects
of a fence configuration and leading-edge roughness
were Investigated at a Reynolds number of 4,000,000.

The low-speed static longitudinal stability character-
istics of a wing having 45o serepback of the quarter-
chord line, an aspect ratio of 8, a taper ratio of 0.45,
and NACA 631A012 airfoil sections parallel to the air
stream were determined in the Langley 19-foot press.
sure tunnel at Reynolds numbers from 1. 5 x 106 to
4. 8 x 106. The effects of combinations of leading-
edge and trailing-edge flape of various spans, upper-
enrface flow-control fences, and a fuselage on the
longitudinal stability were investigated.

Contains results and discussion of a preliminary
wind-tunnel investigation at low speeds of at thin
delta wing equipped with a slotted flap. Results
indicated that the maxidmum lift coefficient could be
increased from 1. 45 to 1. 86 by deflecting the double
slotted flap 50o. The angle of attack of the model
necessary to obtain a given lft coefficient was con-
siderably reduced by the addition of the plotted flaps.

NACA RM L52C11

LOW-SPEED LONGITUDINAL AERODYNAMIC CHAR-
ACTERISTICS OF A TWISTED AND CAMBERED
WING OF 450 FUWEEPBACK AND ASPECT RATIO 8
WITH AND WITHOUT HIGII-LIFT AND 5TALL-
CONTROL DEVICES AND A FUSELAGE AT
REYNOLDS NUMBERS FROM 1.5 x 106 TO 4.8 x 106.
Relnn J. Salmr. June 1952. 76p. diagrs., photo.,
2 tabs. (NACA RM L52Cll)

The low-speed static longitudinal stability character-
istics of a twisted and cambered wing having 450
sweepback of the quarter-chord line, an aspect ratio
of 8, a taper ratio of 0.45, and an NACA 631A012
thickness distribution parallel to the air stream,
were Investigated in the Langley 19-foot pressure
tunnel at Reynolds numbers from 1.5 x 100 to
4.8 x 106. The effects of combinations of leading-
edge and trailing-edge flaps of various spans, upper-
surface flow-control fences, and a fuselage on the
longitudinal stability were investigated.

An investigation was made in the Langley full-scale
tunnel to determine the effects of symmetrically
located wing nacelles on the low-speed aerodynamic
characteristics of a 600 sweptback delta-wing -
fuselage combination. The wing section was an
NACA 65A003 airfoil section and the aspect ratio was
2.31. The model was tested with nacelles at three
chordwlse locations at each of three spanwise sta-
tions. Lift, pitching moment, and drag were ob-
lained through the angle-of-attack range from -3o to
tall at zgro yawr at Re iode numbers from
1.55 x 10o to 2.77 x 10e and Mach numbers from
0.07 to 0.12.

COMPREBSIBLE-FLOW SOLUTIONS FOR THE
ACTUATOR DISK. James B. Delano and John L.
Crigler. March 1953. 70p. diagrs. (NACA
RM L53A07)

Solutions for the actuator disk in subsonic compress-
ible flow are presented. The induced flow phenom-
ena for compreseable and incompressible flow are
shown to be widely different. Solutions have been
obtained for flows with subsonic and supersonic
wakes. Calculations are presented to show the large
gain in efficiency that can be obtained by the use of
disks in tandem for power loadings greater than
those required to choke single disks.

A two-di~mensional investigation of the NACA 64-006,
64-008, 64-010, and 641-012 airfoil sections has
been made In the Langley low-turbulence pressure
tunnel at angles of attack of -2o to 310 and Mlach
numbers of 0.39 to that for tunnel choke. Mdeasure-
ments were made of the lift, drag, and pitching-
moment coefficients for the airfoil models in the
smooth condition and with leading-edge roughness.
One airfoil model, the NACA 641-012, was also
tested with a roughness strip at the 20-percent-
chord station.

A helicopter rotor powered by tip-located ram-jet
engines has been investigated on the Langley
helicopter test tower. The propulsive and aerody-
namic characteristics of the isolated engines were
studied also in the nonwhirling condition in a small
wind tunnel. The basic hovering characteristics,
as well as the aerodynamic and propulsive
characteristics of the engines, have been obtained.

The low~-speed longitudinal stabritty characteristics
of a 450 sweptback wing of aspect ratio 8 having
twist and camlbered alrfolt sections were Investigat-
ed by means of force and pressure-dlstribu tion
measurements at Reynolds numbers froml 1. 5 x 106
to 4 8 x 106 In the Langley 19-foot pressure tunnel.
The effects of Reynolds number, leadrng-edee rough-
ness, upper-surface fences, and leadllna-edge ana
tralling-edge flaps have been determined. A com-
parison has been made between the results obtained
on the twisted and cambered wing and the results
obtained on a wing of similar plan form without twist
and having symmetrical airloil sections of the sa~e
thickness drstributions. The experimental pressure**
dlstrlbutlon loadings have been compared to calcu-
lated loadings.

NACA RM L52K26

THE LOW-SPEED LIFT AND P~ITCHING-MOMENT
CHARACTERISTICS OF A 45o SWEPTBACK WING
OF ASPECT RATIO 8 WITH AND WITHOUT HIGH-
LIFT AND STALL-CONTROL DEVICES AS DETER-
MINED FROM PRESSURE DISTRIBUTIONS AT A
REYNOLDS NUMBER OF 4.0 x 100. Thomas V.
Bollech and William M. Hadaway. January 1953.
57p. diagrs., photo. (NACA RM L52K26)

This paper presents the low-speea dift and pitching-
moment characteristics of a 450 sweptback wing hav-
ing an aspect ratio of 8, a taper ratio of 0. 45, and
incorporating NACA 631A012 airfoil sections in the
streamwrise direction with and without high-lift and
stall-control devices as determined by pressure dis-
tribution at a Reynolds number of 4. 0 x 106 and
through an angle-of-attack range from -40 through
the stall.

A theoretical analysis was made of the centlrlfugal
effects on the fuel-spray paths In a rotating ram-jet
engine tip-mounted on a helicopter blade The
differential equations of motion of the particles were
set up and solved numerically for satly-four selected
combinations of the parameters. Plots of all the
calculated paths are given. A simple method for
determamrng the approximate final direction of the
particles was developed. The results indicated that,
because of the centrifugal action, the larger fuel
particles were likely to hit the outer walls of the
ram jet before burning, so that loss of thrust would
result. Several methods of correcting for this
effect are suggested.

NACA
RESEARCH ABSTRACTS NO.55

The aeroelastic instability of rigid open and closed
bodies of revolution mounted on thin, flexible struts
has been Investigated experimentally at low speeds.
Three types of instability were observed coupled
flutter, divergence, and an uncoupled ascillatory
instability which consults in continuous or intermit-
tent small-amplitude yawang ascillatzons. An
attempt has been made to calculate the airspeeds
and, in the case of the oscillatory phenomena, the
frequencies at which these types of instability occur
by using elender-body theory for the aerodynamic
forces on the bodies.

The effect of centrifugal loadings on the performance
of a helicopter-type ram-jet engine has been de-
termined on the Langley helicopter terit tower.
Halving the centrifugal loading by doubling the radius
reduced the minimum specific fuel consumption at
all ram-jet velocities and Increased the maximum
propulsive thrust by approximately 12.5 percent at a
ram-jet-engine velocity of 630 fps.

An investigation of the effect of leading-edge separa-
tion on the loads exqperienced by a 450 sweptback-
wing model In gusts Indscated that separation in-
creased the load, the amount of load Increase
apparently depending upon the gust-gradient distance
and velocity. It was further indicated that the rate
of change In angle of attack due to the gust and the
extent of po t~etratlon into the gust are both Important
In determining whether separation occurs.

THE FOLLOWING REPORTS HAVE BEEN
DECLASSIFIED FROM RESTRICTED TO
UNCLASSIFIED, 12, 8, 53.

Combustion-chamber performance characteristics
of a 3000-pound-thrust axrial-flow turbojet engine
have been determined from an Investigation of the
complete engine in the Cleveland altitude wind tun-
nel over a range of simulated altitudes and flight

I5

Mach numbers. The effect of variations in altitude
and flight Mach number on combustion efficiency,
combustion-chamber total-pressure losses, engine-
cycle efficiency, and the fractional loss in engine-
cycle efficiency resulting from combustion-chamber
pressure losses is presented for various engine con-
figurations.

Performance of three types of combustion chamber
operating in a 4000-pound-thrust axial-flow turbojet
engine is presented for altitudes from 5000 to
40, 000 feet and ram pressure ratios from 1.00 to
1.86. The loss an cycle efficiency due to the pres-
sure losses in the combustion chamber was found to
be of Little consequence rn the design operating
range of the engine. Combustion efficiency im-
proved with engine speed and ram pressure ratio at
all altitudes and decreased with mecreasing altitude.
At rated engine speed, the altitude effect on combus-
tion efficiency was no greater than 5 percent.
NACA RM E9122

An investigation was conducted in a one-sixth seg-
ment of an annular turbojet combustor to determine
a method of controlling radial exhaust gas tempera-
lure distribution. Two general methods were
studied: (1) ducting the dilution air into the com-
bustion chamber in a predetermined manner through
hollow radial struts and (2) modifying the combustor
basket-wall open-hole area. Results of the rnvesti-
gatlon indicated that: Modificatzons of the hollow
radial struts had some effect on the exhaust-gas
temperature distribution but for the comlbustor in-
vestigated complete control was impossible with this
method. Secondary-zone baskeet-wall modifications,
however, have a large effect on exhaust-gas tem-
perature distribution provided the primary-zone
basket wall Is suitable.
NACA RM L50C23

Model investigations to determine the ditching
characteristics of the Navy XP2V-1 airplane are
described. Various landing configurations were
simulated and the performance of the model was
determined from visual observations, motion-picture
records, and time-history accelerometer records.

Compares three related sets of blades, one con-
ventional in design, one sweptback, and one em-
bodying relatively thin sections. Actual design con-
ditions could not be duphcated in the wind tunnel.
The sweptback blades were in general inferior to the
other two designs; the greatest difference in enve-
Lope efficiency was 2.5 percent. The contributing
factors to the inferiority of the sweptback blades
were Insufficiently high Mach number to show the
full effects of sweep and propeller operation away
from the design conditions.

THE FOLLOWING REPORTS HAVE BEEN
DECLASSIFIED FROM RESTRICTED TO
UNCLASSIFIED, 12,'11,/53.

Methods of introducing and distributing air and fuel
in turbojet-engine combustors were evaluated with
two fuels, AN-F-32 andAN-F-58. Investigations
were made with two single-annulus liners in a one-
quarter sector of a 25-1, 2 inch diameter turbojet

both single-annulus combustors were considerably
higher than those of the double-annulus combustor;
combustion efficiencies were insensitive to changes
in fuel-air ratio.

NACA RM E52117

EFFECT OF MAL;NITUIDE OF VIBRATORY LOAD
SUPERIMPOSED ON MEAN TENSILE LOAD OF
MECHANISM OF AND TIME TO FRACTURE OF
SPEC[MENS AND CORRELATION TO ENGINE
BLADE. Robert R. Ferguson. November 1952.
26p. dlagrs photos., 4 tabs. (NACA RM E52117)

Tensile fatigue tests were run on specimens of seven
turbine blade alloys at a test temperature of 15000 F
and a mean stress of 22, 000 pounds per square mnch
with superimposed alternating stresses of 0, +-5000,
.-10, 000, and '_15, 000 pounds per square Inch. The
same three types of fracture occurring in turbine
blades stress rupture, stress rupture followed by
fatigue, and fatigue were obtained in the specimens.
The type of Iracture obtained was found to be a func-
tlon of the material and the magnitude of the alter-

The results of the investigation indicate that the air-
plane should be ditched at the normal landing atti-
tude with the flaps fully extended. Extensive
damage to the fuselage wil occur and the airplane
probably will dive. II a trapezoidal hydroflap 4 feet
by 2 feet by 1 foot is attached to the airplane at
station 192.4, diving will be prevented.

An experimental investigation was made in the
Langley stability tunnel to determine the effects of
various chordwise fences on the longitudinal sta-
biliry characteristics of an airplane model with a
350 sweptback ~wing. The investigation included the
determination of the effects of fence shape, size'
and position on the longitudinal characteristics of
several model configurations.

Results are presented of an experimental investiga-
tion of the effects of wing-tip gun turrets with vart-
ous modifications upon the aerodynamic character-
isties of a typical bomber-wing model. The data
indicate that the addition of the turrets had negligible
effects upon the lift and pitching-moment character-
istles of the wing. The drag coefficient was in-
creased by 0.005 up to a Mach number of 0.70, and
the Mach number of drag divergence was decreased
by 0.05.

NACA
RESEARCH ABSTRACTS NO.55 1

nating stress. With increasing alternating stress,
the mechanism of failure changed from stress rup-
ture to stress rupture followed by fatigue and then
to fatigue. Six of the seven alloys showed the same
mechanism of failure for specimens tested at an
alternating stress of r5000 pounds per square inch
as for blades tested in a J33-9 let engine.

A general analysis is presented which allows the
rolling motions of an aircraft using a displacement-
response, flicker-type automatic pilot to be deter-
mined, and charts are included for finding the ampli-
tude and period of steady-state oscillations of any
aircraft. Current trends in pilotless-arrcraft de-
signs indicate that small amplitude residual oscilla-
tions are possible with the topic system. The
analysis shows close agreement with roll-simulator
tests.